US2940032A - Instrument landing system - Google Patents

Description

June 7, 1960 A. PROTZEN INSTRUMENT LANDING SYSTEM 3 Sheets-Sheet 1 Filed Aug. 24, 1953 m/nwrop among Profile/z.

June 7, 1960 A. PROTZEN 2,940,032

' INSTRUMENT LANDING SYSTEM Filed Aug. 24, 1953 3 Sheets-Sheet 2 June 7, 1960 A. PROTZEN 2,940,032

INSTRUMENT LANDING SYSTEM Filed Aug. 24, 1953 3 Sheets-Sheet 3 INVENTQP fan 0 7 razzex larly an aircraft, may

: re in ia s l wer s lending tee s a ,Py mean of h c t e p l is I he invention relates .10 a ne and .apgar atus tor d rivi'ngaeomr i less r U aveni iejp'artie I guid on ,toianflele fically p cp 'a ed g i i g eam s ndQD IyPmTthe'indieation bf an 'inst'r'ument'twhichi p'onsi" tothedeyiaitiion frorr'ithe'fguidingbea'fiiQ ltisjtolh i nderstqodthat a co'ntrol effect accordingto thein entio may he used for guiding a yehicleinia' dire remand not onlythe QIiISe thereof infthe hor' aliislan'e'. fiince thefinvem aircrafnih f i lser ption will refeifl'for p 7 aircraft; i Howeyer, it ,Will .be aggarent, .thetgtheinvsil ibn is applicable iin 'generall Flight 'navigations devices, such as airway beacons efi'l ii wnic la lc s'l lfi u a'bled to identify the position bf'the aircraft in relationctda nrescriheiipath.

the airway beacon js enIployic d rfq r overland navigation, the landing beacon is employed for ap A landing. 130th types 6f bieaco adapted for.

ev is W s l .a i st s li er e b t ;1 ,s fti ly on y wi h the lan in beacon. :Yertica area -m i s pa i .t s ef? 1 "t e lsn'di approach.

lnithis type of operations the pilot obtains on an instrument a continuous indication of the distance of ,the aircraft am the VeIti t si e Rl s te me by the guiding beam ,gehejrated hy itlrelbeacqn. For 'ex arnplefin the approach, the riilot has to guide the aircraft'lon m'a vertical rlefiei-enc e plane bisecting the runway] as well as on to a plane through the desired landing point and correspqpdipg to the prescribed gilide jathf In vgl erhl', it is'iiot fl fiidi "c eiefe r'red to, but some other'efiect representing the ation of the hicle Ifr' mt 's ns b i' li n sfiss "shi e-fi 81 9 ans t e s i m bean v i l l lQ n th aircraft to the transmitter that is used to represent the distance. For simplicitygjn the following description of yarious embodiments of t ,e invention reference will be inade to the'deviation'."

' It is a simple matter fertile pilot to determine from indication of "the deviatibn by" an indication, such as a. null indication, which would be easier for -the pilot to 'in'terpre tl' Such deviceslallow the aircrafit to be autom fl fllr ui edcn t es g beam i th p ntmet Tlvmake c n s as *9 I less. x tinserrer 'zer ElY as-sn th b amiis tlnrebyfiscted Without 2,940,032 Patent ed June 7, 1960 ,any mental efiort onstheipart ofthep'ilot. on'jthe'basis tof the readinggofsthe null indicatortlthe pilot, needs only to, I carry. nutssimple.v course or --inclination .corrections in the usual way.

;Iny=order ttosfacilitate; the understanding, of the present linvention theabasic {construction ofssuch a device for-use .With a normal blind landing system (for .example an LL S. System) .twilhnow .beidescribed: with reference to :Fig. $1 of therdrawings. 10

lmentimithoutjintegratipn,

Fig. 4, is a schematic :diagram of .au'devicesimilar to :that .OfiFig. 2 inwhichithe control etfectcan becombined with its :integral,

Fig. 5 is ascircuit-idia'gramiof :an'arrangement similar to thatqof :Fig. 3,; in whieha single null indicator may vhaverselectively appliedttheretothe deviation effect or to one s killed in the art,

sense.

. such an .extentthat "the Output :Identical oriizsim ilar components in'the various figures are indicated .-b.y.lthetsame reference numerals. v ,InaFig. 11, ran indicating instrument 2 indicates to the .pilot intthe .nsual mannentheTdeviation determined by a receiver 1'. Iris-assumed that'the guide beamis vertical, so that the corrections-apply:toa-the course in'the strict The receiver output is. applied not only to the iindicating instrurnentnhut also toa combining amplifier 3, in whichjit istcdmbined-with a voltage de'rivedfrom' a fc'ourseiresponsivedevice14; The" combined amplifier out-,

"The null indigeneral of a sensitive'moving-coil instrument. iThe" course-responsive device comprisesa course-indicating pointer 6, 'which is actuated usually by a master cor'npass (not shown in the drawing), as well as a,col'lrseadjuster 7, which is manually adjusted by the pilot in accordance with the known =I .-L.S. course 'uiion a fixed course scale 8, the arrangement being'such that .the'ihdicating' pointer 6 taps oiT from a potentiometer'9, which iscombined with the course adjuster'7, a'vo ltage proportional to the difierence between the actual course and .the l.L.S. cdiirse to which the course adjusteris set. deviation is then compared with the lateral deviation value in thecombining amplifierS. i

"Jf'th'e aircraft is' now parallel to the set I.L.S. course .the combining amplifier 3 receives'only the deviation -error from the receiver, whilst the output of the coiirse responsive device is zero. "The null'indicator 5 therefore .defiec'ted'until the pilot hascharig'ed his coursfto voltage of'lthe device 4 compensates for the deviation voltage'rec'eive'd froni'the receiver 1. flhe aircraft has thus "assumed a new course directed towards the guiding bea'm'," whereby {reduction cf the deviation is obtained. "However; the courseresponsiye voltage now dominates the combining amplifier 3 and makes-it' necessary for the 'pilot "toreduce the angle of the course with respeettd the guiding beam in Drdertp return the hullindicator again t6 z'erol" This process is repeated until the aircraft has again reached the guiding beam alonga substantiallylogarithmiccurve. Thedeviation voltageandthe'cOfiiSe voltage are then both down to zero "in the amplifier fi 'andthe aircraft thus fliestru on the-beam? By means ofl the above-described method it is thus possible toguidethe aircraftalong a substantially ldgarithrrfic lpath, whichapproaches 'th guiding'beain asyntntotically and to hold thiscour's'e, the pilot having cirn'ly to makecorrections according'tdthe hull arem-i F da. mentally, the process described above a; horizontal orientation-will be a'plilied-ifi a similar' manner tbc riefit'ae v the comparison. approximated by the turning speed of the aircraft about vtion in the vertical plane.

sive device 4 will be replaced by an inclination-responsive device (gyroshorizon). V

' Various modifications of the described-arrangement are known but do not affect the principleof the method. For example, it is known in order to facilitate making corrections according to the null indicator 5' by applying to the combining amplifier 3 an additional effect derived from banking of the aircraft, this effect being derived from a gyro horizon. V

Although the above process makes the blind approach and the staying on the guiding beamsimpler, it is attended by certain drawbacks. Unavoidabl'e errors in the air-.

craft compass installation'may cause erroneoussetting of the course-responsive device 4 and this may give rise to a constant course 'errorin the approach." The effect I that case, the course-respon- I tional to the output voltage of the amplifier, and the of side wind will also have similar results. i The error due to side wind may be avoided by suitable adjustment of the course-responsive device 4, but this presupposes extended "observation and no variation of the side wind component. Since however, due to the constantly changmg altitude, a change of wind must also be accounted for, these c'orrections'are more or less illusory. Owing to these drawbacks it may be impossiblej'under-unfavourable conditions'to make the approach with the desired degree of accuracy, so that the-method often leads 'to unsuccessful approaches on narrow runways,

It is, anobject of. the invention to' provide a-m'ethod and an apparatus for deriving a;control effect for guiding a vehicle -on tov a guiding. beam, which will avoid 'the drawbacks ofprior. methodsand arrangements. Aspecific obje'ct'iof'theinvention'is to provide a method .and an apparatus than-Will :be. independent oferrors in the course setting and/or side wind fluctuations 1 A furtherobject of the inve'ntiomis' .to; reduce the 4 sive device 14. The speed of the motor 15 is propormotor actuates a sliding contact through a gear 16, which sliding contact taps off a proportional voltage on a potentiometer 17. This voltage is applied to a second input of the amplifier 3, the arrangement being such that the voltage tapped from the, potentiometer by the sliding contact is alway proportional tothe deviation effect produced by the receiver.

Ifthis deviation effect varies, the sliding contact moves at a rate proportional to the rate of variation ofthe deviation efiect. This means that the output voltage of theamplifier 3 which is applied to the motor 15 is proportional to the lateral speed of the aircraft, or, what amounts to the samething, proportional to the first derivative ofthe deviation.

A gyro apparatus18 of known construction, which in the described example. is a turn indicator, measures the.

taps off on the potentiometer 20 a voltage proportional to the turning. speed. This turning speed very closely.

approximates the lateral acceleration of the aircraft, that is, therequired second derivative of the deviation. In

' the gyro apparatus 18 the precession moment of the gyro is compared with. the moment produced by an electrical moment generator 21, which is shown in' the present' examplein the form of ajmovable coil, For'thispur ose, the precession axle of the'gyro apparatus is opera- 7 tively connectediwith the armature of an electromagnet. Thederivative efiect orvoltage measuredat the potentio'm eter 17 and the deviation delivered to the motor from the outputoffthe-amplifier are applied to the windyinfluenceof' irregularities or bends" 'in'the guiding beam.

- According to the invention the first derivative with respectto-time is derived from the deviation value and is compared with the deviation. itself.;. In'this manner the guiding'Qf the aircraft:- on to the guiding-"beam is made completely independent of the course, so'that the associated errors, in particular-that 'due to deviation errors of the gyro-compass, are completely avoided. In additionthe influence of side wind fluctuations, which are in effect equivalent with a course'error', is also automatically compensated for in this manner withoutthe pilot having to'concern himself with it. The required course correction is automatically produced and adjusted without even coming to the knowledge of the pilot.

Preferably, an effect representing the second derivative of the deviation with respect to time is introduced into This second derivative is very closely an axis normal to the plane of travel. Whilst therefore the first derivative is derived from thedeviation, the second derivative is preferably measured directly. If the aircraft is then guided by means of the null indicator in ings 1 27 and 2 3' of the movableicoil and there produce .m omen ts',whichj are compared directly with theyprecessionmoment of thegyro, "thatis to' say with the second derivative. The voltage tapped off onthe potentiometer ,29 is therefore proportionalsto the sum of the deviation 7 'volta'ge'and the first and second derivatives, The control 'efiectthus derived applied to an integrating motor 24,

' voltage, which thus represents the integrated control effect, is then applied tothe moving-coil instrument formsuch a way that the control effect formed by the sum of the three values, that is to say of the deviation and of the first and second derivatives thereof, is always zero, then the aircraft is guided on to the beam along a logarithmic curve asymptotic to the beam.

The use of the gyro apparatus for introducing into the comparison the directly measured second derivative increases the accuracy of the method.

7 pilot is familiar.

Further improvements and preferred embodiments of a the inventionwill now be described with reference to Figures 2 to 5 of the accompanying drawings;

7 In the embodiment represented in Fig. 2, -1 represents theJLhS. receiver, the output 'of,which-represents the deviation, and 2 represents the conventional indicating device. 'Besides being applied to theindicator 2, the output of the receiver 1 is also'applied to an input circuit of an amplifier 3. The; amplifier output drives an electric-motor 15 which is ,compiisedin acours'e-responwhich in-turn takes oif a voltage on apotentiometer 26 by means of a slider operated through a gear 25. "This ing the null indicator 27 andis there indicated.

In the modified embodiment according to Fig. 3, the final integration of the control effect is dispensed with and the control effect is applied directly to the null indicator.

If the pilot maintains the null indicator in the zero position by suitable operationof the flight trimming controls, the sum, of the deviation efiect and of the first and second derivatives is always zero and the condition for an asymptotic approach to the guiding beam is fulfilled. The deflections of the null indicator represent,

in the case of Fig. 2, course errors or, in the case of Fig. 3, turning speed errors and consequentlyv can be compensated for easily by the pilot with the aid of familiar techniques. 'In the case of Fig.3 this means that flight according to thenull indicator corresponds exactly to flight according to a turn indicator, with whichevery The integration of the control effect has'the advantage 'that a verysteady indication is obtained. By providing stops 28 for the slider the indicated course error can be limited at a chosen threshold level. Thishas the advantage that at the beginning of theapproach no ,extensive course, changesneed to be e'ifected in order. to bring the indicator to zero. a a p I In the, Fig. 4 embodiment, the. null indicator 27 'provides'a combined indicationof the turning speed error a derived by the gyro apparatus 18 and the course error 1 derived; by theintegrating motor '24; adjustable potentiometer may be provided in the inte- In addition, an

grating circuit for varying the combining ratio. The combination of the course error and the turning speed error makes it much simpler for the pilot to guide the aircraft than if only the course error indication is employed. By means of this combining process the best compromise can be obtained between guiding accuracy and indicator stability.

The Fig. embodiment differs from that of Fig. 3 in that the second null indicator 27 is omitted, the first null indicator 2 being selectively connectable to the gyro device 18 or to the receiver.

It has been mentioned above that in prior methods an effect representative of the banking position Was employed for facilitating the guiding operation. This effect may be produced in a simple manner by the present apparatus by tilting the gyro device 18 in such a manner that the gyro, besides measuring the turning speed of the aircraft about the vertical axis, also measures a turning speed component about the aircraft longitudinal axis the effect thus obtained being, of course, proportional to the inclination at which the gyro is set. By varying this inclination the responsiveness of the gyro to banking can therefore be adjusted as required. The measured banking rate effect is integrated by the integrating device and is applied to the null indicator as a banking position effect.

What I claim is:

1. In an apparatus for deriving a control effect for guiding a vehicle onto an electrically propagated guiding beam, a measuring instrument for deriving a first effect representing the deviation of the vehicle from the guiding beam, means for deriving the first derivative of said effect with respect to time, gyro means responsive to the turning speed of the vehicle about an axis normal to the plane of travel thereof for deriving the second derivative of said effect with respect to time and means for comparing the said derivatives with the said effect to derive a control effect.

2. In an apparatus for deriving a control effect for guiding a vehicle on to an electrically propagated guiding beam, a measuring instrument for deriving an effect representing the deviation of the vehicle from the guiding beam, means for deriving the first derivative of said effect with respect to time, gyro means responsive to the turning speed of the vehicle about an axis normal to the plane of travel thereof for deriving the second derivative of said effect with respect to time, means for comparing the said derivatives with the said effect to derive a control effect, an indicating instrument for giving an indication of the com- '1 bined effect of said effect andsaid first and second derivatives of said effect.

3.Apparatus according to claim 1, for guiding a vehicle onto a vertical guiding beam, in which the said gyro means is adjustable so as to be made responsive to an adjustable degree to the turning of the vehicle about its longitudinal axis for deriving a second effect representing the turning speed of the vehicle about the said longitudinal axis, the said second effect being applied to a null indicator.

4. In an apparatus for deriving a control effect for guiding a vehicle on to an electrically propagated guiding beam, a measuring instrument for deriving an effect representing the deviation of the vehicle from the guiding beam, means for deriving the first derivative of said effect with respect to time, gyro means responsive to the turning speed of the vehicle about an axis normal to the plane of travel thereof for deriving the second derivative of said effect with respect to time, means for comparing the said derivatives with the said effect to derive a control effect, and an integrator having the said control effect applicable thereto for deriving therefrom the integrated value thereof.

5. Apparatus as claimed in claim 4, comprising an indicating instrument for indicating the said integrated value, said instrument having adjustable stops for limiting the value indicated thereby when the said integrated value surpasses a predetermined threshold level.

6. Apparatus as claimed in claim 4, comprising a null indicator to which the said control effect and the integrated value thereof is applicable in predetermined proportions.

Rcfereuces Cited in the file of this patent UNITED STATES PATENTS 1,703,280 Minorsky Feb. 26, 1929 1,958,258 Alexanderson May 8, 1934 2,259,600 Alkan Oct. 21, 1941 2,372,185 Wittkuhns Mar. 27, 1945 2,423,337 Moseley July 1, 1947 2,513,537 Williams July 4, 1950 2,548,278 Wirkler Apr. 10, 1951 2,613,350 Kellogg Oct. 7, 1952 2,613,352 Kellogg Oct. 7, 1952 2,644,941 Kellogg July 7, 1953 2,759,137 Kutzler Apr. 14, 1956 OTHER REFERENCES Aero Digest, January 1949, pages 58, 59 and 87.

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